The main structure of the fuselage of an aircraft typically comprises a skin with stringers and frames. The skin is stiffened longitudinally by stringers to reduce its thickness and be competitive in weight, while the frames prevent the general instability of the fuselage and may be subjected to local load inputs.
As the aeronautical industry requires structures which, on the one hand, must support the loads to which they are subjected, thus fulfilling high requirements of strength and stiffness, and, on the other hand, must be as light as possible, the use of composite materials in aircraft primary structures has been increasingly spread since by conveniently using said composite materials significant weight savings can be achieved compared to traditional designs made of metallic materials.
In areas of aircraft structures with openings or subjected to high loads, for example, those areas of the fuselage of an aircraft withstanding the load inputs from the floor of the aircraft cabin, special requirements arise.
FIGS. 1a and 1b show two known solutions in the art to meet these requirements in the case of fuselages which use longitudinal beams instead of stringers in areas with openings to stiffen and reinforce them. Crosses between beams and frames are made such that only one of the two elements is maintained continuous (whether the beams or the frames). This implies that both elements shall be thereafter joined in the crossing zones by a plurality of connecting elements (riveted or bonded), resulting in high weight penalties and/or high production and assembly costs and also debonding risks when using connecting elements bonded to the beams and the frames.
The present invention is directed to solving these problems.